U.S. patent number 5,702,232 [Application Number 08/354,299] was granted by the patent office on 1997-12-30 for cooled airfoils for a gas turbine engine.
This patent grant is currently assigned to United Technologies Corporation. Invention is credited to Robert P. Moore.
United States Patent |
5,702,232 |
Moore |
December 30, 1997 |
Cooled airfoils for a gas turbine engine
Abstract
An internally cooled airfoil for a gas turbine engine includes a
leading edge region, a mid-chord region, a transition region, and a
trailing edge region. The mid-chord region has a double wall
configuration with the two outer walls and two inner walls. The
trailing edge region has a conventional single wall configuration.
The transition region has a three wall configuration designed to
provide a gradual transition from the four wall configuration to
the two wall configuration and to minimize high stresses
therein.
Inventors: |
Moore; Robert P. (Tequesta,
FL) |
Assignee: |
United Technologies Corporation
(Hartford, CT)
|
Family
ID: |
23392694 |
Appl.
No.: |
08/354,299 |
Filed: |
December 13, 1994 |
Current U.S.
Class: |
416/95; 416/96R;
416/97R |
Current CPC
Class: |
F01D
5/186 (20130101); F01D 5/187 (20130101); F05D
2260/202 (20130101) |
Current International
Class: |
F01D
5/18 (20060101); F01D 005/18 () |
Field of
Search: |
;416/9R,91,92,95,96R,97R,97A,96A |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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|
|
1210253 |
|
Feb 1963 |
|
DE |
|
243324 |
|
Sep 1969 |
|
SU |
|
895077 |
|
May 1962 |
|
GB |
|
Primary Examiner: Carone; Michael J.
Assistant Examiner: Wesson; Theresa M.
Claims
I claim:
1. An airfoil for a gas turbine engine having a suction side wall
on a suction side and a pressure side wall on a pressure side, said
pressure side wall and said suction side wall extending from a
leading edge to a trailing edge in a chordwise direction, said
airfoil having a leading edge region, a mid-chord region, and a
trailing edge region, said regions sequentially situated in said
chordwise direction from said leading edge to said trailing edge,
said mid-chord region having at least one suction side feed passage
and at least one pressure side feed passage, said suction side feed
passage being disposed on said suction side and bound by said
suction side wall and a first inner wall, said pressure side feed
passage being disposed between said pressure side wall and a second
inner wall, a feed chamber being defined between said first inner
wall and said second inner wall, said trailing edge region having
at least one trailing edge passage disposed between said pressure
side wall and said suction side wall, said airfoil characterized
by:
a transition region disposed between said mid-chord region and said
trailing edge region, said transition region having a transition
feed passage and a transition chamber, said transition feed passage
and said transition chamber being separated by said first inner
wall extending from said mid-chord region into said transition
region in said chordwise direction and bound by said suction side
wall and said pressure side wall.
2. An airfoil for a gas turbine engine having a suction side wall
on a suction side and a pressure side wall on a pressure side, said
pressure side wall and said suction side wall extending from a
leading edge to a trailing edge in a chordwise direction, said
airfoil having a leading edge region, a mid-chord region, and a
trailing edge region, said regions sequentially situated in said
chordwise direction from said leading edge to said trailing edge,
said mid-chord region having at least one suction side feed passage
and at least one pressure side feed passage, said suction side feed
passage being disposed on said suction side and bound by said
suction side wall and a first inner wall, said pressure side feed
passage being disposed between said pressure side wall and a second
inner wall, a feed chamber being defined between said first inner
wall and said second inner wall, said trailing edge region having
at least one trailing edge passage disposed between said pressure
side wall and said suction side wall, said airfoil characterized
by:
a transition region disposed between said mid-chord region and said
trailing edge region, said transition region including a transition
chamber, said pressure side feed passage extending into said
transition region, said pressure side feed passage extending into
said transition region, said pressure side feed passage and said
transition chamber being separated by said second inner wall
extending into said transition region in said chordwise direction
and bound by said suction side wall and said pressure side wall to
reduce stress concentration in said airfoil.
3. An airfoil for a gas turbine engine having a suction side wall
on a suction side and a pressure side wall on a pressure side, said
pressure side wall and said suction side wall extending from a
leading edge to a trailing edge in a chordwise direction, said
airfoil having a leading edge region, a mid-chord region, and a
trailing edge region, said regions sequentially situated in said
chordwise direction from said leading edge to said trailing edge,
said mid-chord region having at least one suction side feed passage
and at least one pressure side feed passage, said suction side feed
passage being disposed on said suction side and bound by said
suction side wall and a first inner wall, said pressure side feed
passage being disposed between said pressure side wall and a second
inner wall, a feed chamber being defined between said first inner
wall and said second inner wall, said trailing edge region having
at least one trailing edge passage disposed between said pressure
side wall and said suction side wall, said airfoil characterized
by:
a transition region disposed between said mid-chord region and said
trailing edge region, said transition region including a transition
chamber, said suction side feed passage extending into said
transition chamber, said suction side feed passage extending into
said transition region, said suction side feed passage and said
transition chamber being separated by said first inner wall
extending into said transition region in said chordwise direction
and bound by said suction side wall and said pressure side wall to
minimize stress concentration in said airfoil.
4. An airfoil for a gas turbine engine having a suction side wall
on a suction side and a pressure side wall on a pressure side, said
pressure side wall and said suction side wall extending from a
leading edge to a trailing edge in a chordwise direction, said
airfoil having a leading edge region, a mid-chord region, and a
trailing edge region, said regions sequentially situated in said
chordwise direction from said leading edge to said trailing edge,
said mid-chord region having at least one suction side feed passage
and at least one pressure side feed passage, said suction side feed
passage being disposed on said suction side and bound by said
suction side wall and a first inner wall, said pressure side feed
passage being disposed between said pressure side wall and a second
inner wall, a feed chamber being defined between said first inner
wall and said second inner wall, said trailing edge region having
at least one trailing edge passage disposed between said pressure
side wall and said suction side wall, said airfoil characterized
by:
a transition region disposed between said mid-chord region and said
trailing edge region, said transition region having a transition
feed passage and a transition chamber, said transition feed passage
and said transition chamber being separated by said second inner
wall extending from said mid-chord region into said transition
region in said chordwise direction and bound by said suction side
wall and said pressure side wall.
Description
CROSS REFERENCE TO RELATED APPLICATIONS
The subject matter of this application is related to the subject
matter of commonly owned U.S. patent application Ser. No.
07/236,092 filed on Aug. 24, 1988 and entitled "Cooled Blades For A
Gas Turbine Engine" still pending.
TECHNICAL FIELD
This invention relates to gas turbine engines and, more
particularly, to internally cooled airfoils therefor.
BACKGROUND OF THE INVENTION
A gas turbine engine includes a compressor, a combustor, and a
turbine. Air flows axially through the engine. As is well known in
the art, the compressed gases emerging from the compressor are
mixed with fuel in the combustor and burned therein. The hot
products of combustion, emerging from the combustor at high
pressure, enter the turbine where the hot gases produce thrust to
propel the engine and to drive the turbine, which in turn drives
the compressor.
Both the compressor and the turbine include alternating rows of
rotating and stationary airfoils. The airfoils operate in an
especially hostile environment that is characterized by high
pressure, high temperature, and repeated thermal cycling. For
example, the temperature of hot combustion gases entering the
turbine generally exceeds the melting point temperatures of the
alloys from which the turbine airfoils are fabricated. Thus, to
properly perform in such a harsh environment, the airfoils must be
cooled.
One effective cooling method is described in U.S. patent
application Ser. No. 07/236,092 to Auxier et al, entitled "Cooled
Blades For A Gas Turbine Engine" and assigned to Pratt &
Whitney, a division of United Technologies Corporation of Hartford,
Conn., the assignee of the present invention. The disclosed airfoil
includes a double wall configuration in the mid-chord region
thereof with a plurality of radial feed passages defined on each
side of the airfoil between an inner wall and an outer wall. A
central radially extending feed chamber is defined between the two
inner walls. The trailing edge of the airfoil, shown in the
aforementioned patent application, includes a conventional single
wall configuration with two outer walls defining a sequence of
trailing edge passages therebetween.
Although the disclosed airfoil provides an effective cooling
configuration, an improvement is needed to minimize stress in the
interface area between the double wall configuration in the
mid-chord region and the single wall configuration in the trailing
edge. The high local stress in the interface area results from the
transition from a four wall configuration in the mid-chord section
to a two wall configuration in the trailing edge. In addition, a
thermal difference within the interface area contributes to high
stress therein. The thermal difference arises as the two outer
walls in the mid-chord region merge with the two inner walls and
transition to two walls in the trailing edge region. The outer
walls of the mid-chord region are exposed to the hot gases and,
despite the cooling, remain relatively hot. The two inner walls are
shielded by the feed passages on one side and cooled by the feed
chamber on the other side, and therefore remain relatively cool.
The two outside walls in the trailing edge region remain relatively
hot. As the two hot outer walls and two cooler inner walls of the
mid-chord region merge into two hot walls in the trailing edge
region, high stress is produced. These high stresses in the
interface area must be reduced to improve the durability of the
airfoil.
DISCLOSURE OF THE INVENTION
According to the present invention, a gas turbine engine airfoil
having a leading edge region, a mid-chord region, and a trailing
edge region with the mid-chord region having a double wall
configuration and with the trailing edge region having a single
wall configuration includes a transition region between the
mid-chord region and the trailing edge region to minimize stress
within the airfoil. The transition region includes a transition
feed passage and a transition chamber. The transition feed passage
and the transition chamber are separated by a transition inside
wall, thereby resulting in a three wall configuration within the
transition region.
The three wall configuration in the transition region reduces
stress in the airfoil. The stress is minimized due to gradual
structural and thermal transitions from the four wall configuration
in the mid-chord region to the three wall configuration in the
transition region to the two wall configuration in the trailing
edge region. Furthermore, the transition feed passage provides
shielding to the transition chamber, thereby improving cooling in
the transition region. Cooler air in the transition region also
translates into cooler air in the trailing edge, thereby improving
cooling within the trailing edge region.
The foregoing and other advantages of the present invention become
more apparent in light of the following detailed description of the
exemplary embodiments thereof, as illustrated in the accompanying
drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a schematic representation of a gas turbine engine;
FIG. 2 is a perspective view of an airfoil of the gas turbine
engine of FIG. 1;
FIG. 3 is a sectional view of the airfoil of FIG. 2, according to
the present invention;
FIG. 4 is a sectional view of an airfoil according to the prior
art; and
FIG. 5 is a sectional view of an alternate embodiment of the
airfoil of FIG. 2.
BEST MODE FOR CARRYING OUT THE INVENTION
Referring to FIG. 1, a gas turbine engine 10 includes a compressor
12, a combustor 14, and a turbine 16. Air 18 flows axially through
the engine 10. As is well known in the art, air 18 is compressed in
the compressor 12. Subsequently, the compressor air is mixed with
fuel and burned in the combustor 14. The hot products of combustion
enter the turbine 16, wherein the hot gases expand to produce
thrust that propels the engine 10 and drives the turbine 16, which
in turn drives the compressor 12.
Both the compressor 12 and the turbine 16 include alternating rows
of rotating and stationary airfoils 20. Referring to FIG. 2, each
airfoil 20 includes a platform 22 and an airfoil portion 24. The
airfoil portion 24 includes a pressure side 26 and a suction side
28 extending in the chordwise direction from a leading edge 30 to a
trailing edge 32 and in the spanwise direction from a root 34 to a
tip 36. Referring to FIG. 3, each airfoil portion 24 includes a
leading edge region 40, a mid-chord region 42, a transition region
44, and a trailing edge region 46. The regions 40, 42, 44, and 46,
extend sequentially chordwise from the leading edge 30 to the
trailing edge 32.
The mid-chord region 42 includes a double wall configuration having
outer walls 48, 50 and inner walls 52, 54 on the pressure side 26
and suction side 28, respectively. A plurality of pressure side
feed passages 56 is disposed between the outer wall 48 and the
inner wall 52 in the mid-chord region 42. A plurality of suction
side feed passages 57 is disposed between the outer wall 50 and the
inner wall 54 in the mid-chord region 42. A feed chamber 58 is
disposed between the two inner walls 52, 54 in the mid-chord region
42. A plurality of resupply holes 59 allows communication between
the feed passages 56, 57 and the feed chamber 58.
The transition region 44 includes a transition feed passage 60 and
a transition chamber 62 separated by a transition inside wall 64
and bounded by the pressure side outer wall 48 and the suction side
outer wall 50. The cooling air exits the feed passages 56, 57 and
the transition feed passage 60 through a plurality of film holes 65
formed within the outer walls 48, 50.
The trailing edge region 46 includes a plurality of trailing edge
passages 66 sequentially situated therein and bound by the pressure
side outer wall 48 and the suction side outer wall 50.
During operation of the gas turbine engine 10, cooling air enters
the feed passages 56, 57 and feed chamber 58 at the root 34 and
centrifuges towards the tip 36 of the airfoil 20. Cooling air exits
the feed passages 56, 57 through the film holes 65 to cool the
outer walls 48, 50. As the cooling air is depleted through the film
holes 65, the feed passages 56, 57 are resupplied with the cooling
air from the feed chamber 58 through the resupply holes 59.
Although cooling air exiting through the film holes 65 cools the
outer walls 48, 50 in the mid-chord region, the outer walls 48, 50
remain relatively hot. The inner walls 52, 54 in the mid-chord
region are shielded from the outside ambient temperatures by the
feed passages 56, 57 on one side of each wall 52, 54 and cooled by
the cooling air circulating in the feed chamber 58 on the other
side of each wall 52, 54. Therefore, the inner walls 52, 54 remain
relatively cold.
The outside walls 48, 50 in the transition region 44 remain
relatively hot. The transition inside wall 64 is relatively cool
because it is shielded by the transition feed passage 60 on the
pressure side 26 and by the transition chamber 62 on the suction
side 28.
The trailing edge cooling is accomplished by cooling air
circulating through the trailing edge passages 66. The outside
walls 48, 50 on the pressure side 26 and the suction side 28,
respectively, in the trailing edge region 46 remain relatively
hot.
The inclusion of the transition region minimizes stress in the
airfoil 20, over the prior art airfoil 120 as depicted in FIG. 4.
The prior art airfoil 120 does not have a transition region,
resulting in excessive stress at the interface between the
mid-chord region 142 and the trailing edge region 146. The
excessive stress arises due to the double wall configuration
transitioning to a single wall configuration.
One advantage of the airfoil of the present invention over the
airfoil of the prior art is that stress is minimized. The stress is
significantly reduced because the four wall configuration in the
mid-chord region 42 transitions to a three wall configuration in
the transition region 44 and eventually to a two wall configuration
in the trailing edge region 46, as shown in FIG. 3. The transition
region 44 aiso provides for gradual thermal transition of merging
two hot outer walls 48, 50 and two cooler walls 52, 54 of the
mid-chord region with the two hot walls of the trailing edge
region. The gradual structural and thermal transitions decrease the
stress in the airfoil.
Another advantage of the present invention is that the transition
feed passage 60 provides shielding to the transition chamber 62.
The shielding allows the air within the trailing edge passage 62 to
remain cooler and therefore supply cooler air to the trailing edge
passages 66. The configuration of the present invention provides
better cooling in the trailing edge than the airfoil 120 of FIG.
4.
Although the transition feed passage 60 is depicted as being
disposed on the pressure side 26 of the transition chamber 62, the
transition feed passage 60 can be disposed on the suction side 28
of the transition chamber 62. The exact positioning of the
transition feed passage 60 does not alter the benefits of the
present invention of minimizing thermal loading in the transition
region and providing additional shielding and cooling in the
trailing edge region.
An alternate embodiment of the present invention is illustrated in
FIG. 5. The airfoil 220 is analogous to the airfoil 20 of FIG. 3,
with the exception of feed passages 56, 57. A single pressure side
feed passage 256 extends through the mid-chord region 242 of the
airfoil 220 on the pressure side 226 thereof. A single suction side
feed passage 257 extends through the mid-chord region 242 of the
airfoil 220 on the suction side 228 thereof. The suction side feed
passage 257 extends into the transition region 244 of the airfoil
220 to provide shielding for the transition chamber 262 and to
provide a gradual transition from a four wall configuration in the
mid-chord region 242 to a two wall configuration in the trailing
edge region 246.
Although the invention has been shown and described with respect to
exemplary embodiments thereof, it should be understood by those
skilled in the art that various changes, omissions, and additions
may be made thereto, without departing from the spirit and scope of
the invention. For example, the feed chamber 58 may include a
plurality of smaller feed chambers.
* * * * *